The present invention relates to improvements made to gas turbine blades for an airplane engine. More particularly, the invention relates to the cooling circuits of such blades.
It is known that moving blades of airplane gas turbine engines, in particular blades of high pressure turbines, are subjected to very high temperatures from combustion gases when the engine is in operation. These temperatures reach values that are well above those that can be withstood damage by the various pieces that come into contact with these gases, thereby limiting the lifetime of such pieces.
Furthermore, it is known that raising the temperature of the gases in a high pressure turbine improves the efficiency of an engine, and thus the ratio of engine thrust over the weight of an airplane propelled by the engine. Consequently, efforts are made so as to provide turbine blades that are capable of withstanding higher and higher temperatures.
In order to solve this problem, it is known to provide such blades with cooling circuits seeking to reduce the temperature of the blades. By means of such circuits, cooling air, generally inserted into the blade via its root, passes through the blade following a path defined by cavities formed inside the blade prior to being ejected via orifices opening through the surface of the blade.
However, it is often found that the heat exchange produced by this flow of cooling air in the cavities of the blade is not uniform and gives rise to temperature gradients that penalize blade lifetime.
Furthermore, exhausting cooling air through outlet positions in the convex face is difficult. The speeds that apply over the convex face of the blade are high so that the losses that result from mixing between the cooling air and the air in the external stream are high and spoil the efficiency of the gas turbine.